Transiting from Air to Space
The North American X-15

SECTION II
DESIGNING FOR MACH 6

Although the invitation-to-bid letter circulated to prospective contractors by the Air Materiel Command had specifically excluded the NACA Preliminary Study as a requirement, North American's winning proposal bore an unsurprising resemblance to the design envisioned by that study. A comparison of the suggested configuration contained in the NACA study and the North American configuration presented to the first industry conference in October 1956 revealed that the span of the X-15 had been reduced from the 27.4 feet of the suggested configuration to only 22 feet and that the North American fuselage had grown from the suggested 47.5 foot overall-length to 49 feet. The North American design contained 'the split tail surfaces, the wing and tail flaps, the leading-edge sweep for both wing and tail surfaces, and the skid-type landing gear which had been suggested by the preliminary study. The all-movable tail of the 1956 configuration still retained the thick wedge airfoil envisioned by NACA and the horizontal tail surfaces incorporated the cathedral (downward slope or negative dihedral) which had also been a feature of the NACA suggestions. The major differences in external configuration between the study proposal and the design which North American presented consisted of an elimination of ailerons and of separate stabilizers and elevators. North American eliminated the ailerons and elevators by utilizing all-movable horizontal tail surfaces that could be operated differentially so as to provide roll as well as pitch control (the "rolling tail"). North American had gained considerable experience with all-movable controls through using them on the F-107 fighter design, and in this instance use of the differentially operated surfaces permitted simplification of wing construction and elimination of the protuberances that would have been necessary if aileron controls had been incorporated in the thin airfoil sections of the X-15's wings. Such protuberances would have disturbed the airflow and created another heating problem. One other significant difference between the configuration of the NACA design and that of the X-15 stemmed from North American's incorporation of the propellant tanks in the fuselage structure and the use of tunnels on both sides of the fuselage to accommodate the propellant lines and engine controls that ordinarily would have been contained within the fuselage. North American followed the NACA suggestions by selecting Inconel alloy as the major structural material and in the design of a multi spar wing with extensive use of corrugated webs. 1

The original North American proposal gave rise to several questions which in turn, on 24-25 October 1955, prompted a meeting attended by NACA and WADC personnel at Wright-Patterson Air Force Base. The purpose was to consider necessary changes in North American's preliminary design. The meeting formulated a list of questions and comments to serve as the basis of discussions with the contractor. Subsequent meetings of the WADC-NACA group with North American's engineers were held at the contractor's Inglewood plant on 28-29 October and 14-15 November. The items considered at the October and November meetings included North American's use of fuselage tunnels and the rolling tail. The government agencies expressed concern that the tunnels might create undesirable vortices that would interfere with the vertical tail, and suggested that the tunnels be kept as short as possible in the area ahead of the wing. North American agreed to make the investigation of the tunnels' effects a subject of an early inquiry in the model testing program. The contractor also agreed that the "rolling tail" should be proved or disproved as quickly as possible. a

NACA computations indicated that the minimum design dynamic pressure should be 2,100 pounds per square foot and that 2,500 pounds per square foot would be desirable, while North American's design had proposed a design dynamic pressure of only 1,500 pounds per square foot. A structural weight increase of slightly over 100 pounds would enable the design to withstand the 2,500 pound pressure; conferees agreed that the weight increase was justified and that North American should alter the design to meet the 2,500 pound per square foot requirement. On the other hand, a government request that the design be altered to increase the design load factor from 5.25g to 7.33g at a 30-percent fuel-remaining condition involved a weight increase of another 135 pounds which the agencies and North American agreed might better be used to raise the design dynamic pressure. North American also agreed to raise the 35 feet per second negative gust velocity of the design to the 55 feet per second considered desirable by the government representatives.

In addition to the discussions on structural criteria, considerable attention was devoted to the proposed structural materials. At the time of the meeting, neither the WADC-NACA representatives nor the North American engineers seemed to have any detailed information that would permit a final decision on the materials to be used in such critical structures as the leading edge of the wing and the dive brakes. Such diverse materials as plastic, titanium carbide, copper, and cermets (ceramic metallics) were considered for the leading edges; the only definite conclusion was that North American would investigate the relative advantages of several proposed materials. It was agreed to retain the design features which would enable the leading edge to be easily detached and replaced. The NACA-WADC team pointed out that the assumption of laminar flow in heating calculations was unrealistic and North American agreed to build in accordance with the results obtained from calculations based on both laminar and turbulent flow. It was also agreed that 0.020-inch titanium alloy was a more desirable material for the internal structure of the wings and horizontal and vertical stabilizers than the 24ST aluminum that had been proposed, even though the use of titanium produced a weight increase of approximately seven pounds. Another weight increase of 13 pounds was approved in order to allow the substitution of an Inconel-X sandwich construction in place of the stainless steel dive brakes proposed by North American, and to allow for additional dive brake hinges. Other structural problems discussed included a change from titanium to Inconel-X for the oxygen tank because of the low-impact strength of titanium at low temperatures and the need to include a pressure system for stabilizing the propellant tanks. Pressurization of the propellant tanks had been considered undesirable and in the original design had not been provided. The decision to increase the design dynamic pressure from the original 1,500 to 2,500 pounds, together with North American's previous decision to utilize the tanks as structural components, made it necessary to accept pressurization or a large increase in structural weight. The decision was for pressurization of the tanks.

The WADC-NACA group and the North American engineers were in agreement that provision would have to be made for correcting any thrust misalignment and that further investigation would be needed to determine how such misalignment could be corrected and the amount of misalignment that would not be amenable to corrective shimming. The fact that the proposed design would probably be sensitive to roll-yaw coupling was also discussed and the acceptable limits were agreed upon.

In the area of control systems, the WADC-NACA group pointed out to the North American engineers that a rate damping system in pitch and yaw and possibly in roll would probably be necessary. North American estimated that the damping system would increase the weight of the design by approximately 125 pounds. A decision as to whether duplication of the damping system would be necessary was postponed until NACA's Ames Laboratory could be consulted. Conferees also decided that no damping system would be needed in the space control system. It was tentatively agreed that the pilot's controls should consist of a conventional center stick but that the aerodynamic controls should also be operable from a side controller on the right console and that the space controls would be operated by a second side-controller on the left console. The space control system was the subject of further discussion that ended with North American's agreement to duplicate the entire system and to provide three and one-half times the hydrogen peroxide' initially specified. The company also agreed to study the system with a view to minimizing fire hazards, shortening the peroxide lines, and relocating the peroxide supply nearer to the center of the airplane. Separate sources of peroxide would be provided for the reaction controls and the auxiliary power units. Engineers estimated that such changes in the reaction control system would result in a weight increase of about 117 pounds.

At the time of the meeting it was thought that WADC already had a satisfactory stable platform and it was agreed that this platform would be provided as government furnished equipment. NACA promised to provide a nose (then in the development stage) that would contain flight-path indication equipment.

Power-plant discussions were limited because the engine was still subject to extensive development and detailed information was nonexistent. The conference group decided, however, to increase the amount of helium provided for pressurization of the liquid oxygen tank, to study the possible relocation of the helium supply to some area other than inside the oxygen tank, and to redesign the tank transfer tube inlets and the top-off system. Pressure systems were to be protected with relief valves and frangible disks or with duplicate relief valves. The number of engine restarts was to be raised from three to at least five, shut off valves were to be provided in the main propellant lines, and provision was to be made for selective jettisoning of the propellants. Peroxide tanks were to be compartmented and separated, particularly from the engine compartment; main propellant vents were also to be separated and located at the rear end of the jettison lines. Blow-off doors were to be put around the engine compartment and it was agreed to omit a thrust measuring system because of the additional complication such a system would entail.

Final decisions on the exact nature of the auxiliary power plants were delayed to permit further study but there was general agreement that two auxiliary power units should be provided and that they should include completely separate systems compartmented by fire walls.

The discussion between the government and company engineers covered several additional fields including ground check-out equipment, tankage, crew provisions, landing gear, ground equipment, the electrical system, and fire detection and extinguishing. With the exception of the crew provisions, these items were rather briefly considered and included such decisions as the use of nitrogen as the fire extinguishing agent both for ground and air use, the recalculation of the tankage requirements for liquid oxygen, the possibility of providing a jettisonable ventral fin, the various types of ground servicing equipment that would be necessary, the need for providing adequate electrical power for restarts, and.making the electrical components explosion safe.

The discussions on crew provision were more detailed. North American agreed to design an ejection seat system and to make a study justifying the selection of a seat in preference to a capsule system. North American was also to provide suitable head and limb restraints for the accelerations to be encountered and to provide a means for external depressurization and canopy removal independent of the internal canopy jettison system. Transparent cockpit materials were to be studied (transparent plastics, like plexiglass were considered unsatisfactory) and deviations from standard cockpit dimensions were authorized. A gaseous oxygen system replaced the originally proposed liquid system and provisions were included for ram air ventilation below 20,000 feet. Nitrogen was to be used for cockpit pressurization.

The meetings came to a close with a presentation by Douglas engineers of some of the ideas contained in that company's X-15 design and a presentation by North American of its own rocket engine proposals for the X-15. The North American engine would have used oxygen and JP-4 or gasoline as the propellants. North American also presented the results of performance calculations based on the changes that had been discussed and determined upon at the meeting. 2

By January of 1956, North American's design had progressed rapidly enough to require decisions on several questions that had not been discussed or on which no final decisions had been reached at the October-November meetings of the previous year. An NACA group visiting North American on January 18 was asked to provide additional information and guidance on a plan to use a removable instrument rack for the main instrument compartment. Some instruments were to be mounted permanently in the fuselage tunnels, but North American felt a removable rack would provide ready access to the instruments and allow the removal of the instruments during ground operations. This latter feature was considered desirable in order to reduce the exposure of the instruments to ammonia fumes. North American also requested drawings of NACA research instruments and a statement as to which instruments would need to be shock mounted so that the company could complete its instrumentation plans. The company had also reached a stage in the design that required definite decisions on type and gauge of the wire to be used for thermocouples. North American also advised the NACA representatives of plans to use a modified ARC-48 radio communicator with four channels.

The subject of a stable platform came up b and, contrary to the statements made at the October-November meetings of the previous year (that Wright Field had a stable platform and would furnish it to the X-15 project), the NACA group was advised that no decision had been made as to who would furnish the platform. The company asked for further information on instrument duty cycles because without such information, engineers were having difficulty in determining auxiliary power plant loads and heating and cooling requirements.

The NACA representatives agreed with a North American suggestion that the ammonia tank vents could be closed on the ground after filling, thus permitting the pressure to stabilize at the vapor pressure of ammonia. As this procedure prevented boil off of ammonia, it eliminated the necessity for an ammonia top-off system.

Preliminary sketches of the aerodynamic side-controller were shown to the NACA group and as the sketches looked promising, plans were made to have NACA's Langley Laboratory evaluate the system envisaged by North American.

Other topics discussed at this time included the design of the dive brakes and a landing study conducted by North American. Full extension of the dive brakes at pressures of 2,500 pounds per square foot would have created excessive longitudinal accelerations and the brakes were therefore designed to open only to a point where the pressure on them would be 1,500 pounds per square foot. The brakes would then open progressively, maintaining a constant pressure at the 1,500 pound level until the full open position was reached. 3

During the spring and summer of 1956, several scale models were exposed to rather intensive wind tunnel tests. A i/50-scale-model was tested in the 11-inch hypersonic and the 9-inch blowdown tunnels at Langley, and another in a North American tunnel. A i/15-scale model was also tested at Langley and a rotary-derivative model was prepared for test at the Ames Laboratory. North American gave thought to a plan to mount a small model on the nose of a rocket in order to obtain heat-transfer data under flight conditions. Langley, not fully approving of North American's plans, undertook the study of possible alternatives. The various wind tunnel programs included investigations of the speed brakes, horizontal tails without dihedral, several possible locations for the horizontal tail, modifications of the vertical tail, the fuselage side fairings, and control effectiveness. Another subject in which there was considerable interest was that of determining the cross-section radii for the leading-edges of the various surfaces. A free-flight model tested at Langley indicated that the X-15 would have satisfactory handling characteristics. (The NACA studies confirmed the desirability of control system dampers, while during the same period, North American arrived at the conclusion that the airplane could be flown safely without them.)

At the conclusion of a meeting of NACA, WADC, Navy, and North American representatives held at WADC on 2-3 May for the purpose of settling upon specifications, the subject of escape was taken up once more. WADC personnel apparently were not convinced that the ejection seat previously decided upon was going to be adequate, They pointed out that Air Force policy required an enclosed system in all new airplanes and that a change to some form of capsule would not only be in accordance with this policy, but would provide research data on such escape systems. Those opposed to the WADC view objected to any change on the grounds that it would disrupt time schedules, increase weight, and that there was still considerable ignorance about capsule design. The group that opposed the change felt that the safety features of the X-15's structure made the ejection seat acceptable. As a result of this meeting, North American was asked to document the arguments justifying the use of the ejection seat.

A meeting, held at Langley on 24 May and attended by WADC personnel as well as NACA, North American, and Eclipse-Pioneer representatives, explored the possibilities for obtaining a suitable stable platform for the X-15. It appeared that such a platform could be ready in 24 months and that 40 pounds of the estimated total weight of 65 pounds could be charged to research instrumentation rather than to the aircraft itself. 4

By June, NACA had completed the preliminary design for the spherical nose cone and had undertaken the construction of a heat transfer model. They were in the process of preparing detailed specifications for the award of a contract for the cone and its drive mechanism. 5

June was also the month in which formal assignment of Air Force serial numbers was made. The numbers were 56-6670 through 56-6672. Originally furnished by telephone on 28 May, these numbers were officially confirmed by the acting chief of the Contract Reporting and Bailment Branch on 15 June. 6

By July, NACA felt that sufficient progress had been made on the design problems presented by the X-15 to make an industry conference on the project worthwhile. Dr. Dryden, the director of NACA, invited WADC to participate in such a conference and asked that WADC review any material that might be suitable for presentation at the proposed conference. Dr. Dryden also asked that such material be summarized prior to August 8, as that date had been selected for a preliminary meeting of NACA, WADC, Navy, and North American representatives. The participants in the August meeting were to review the summarized material, decide whether the material was of sufficient interest to warrant an industry-wide meeting, and if the material did prove interesting, to make definite plans for a program to be conducted in October at one of the NACA's own facilities. 7

The material did prove interesting and the proposed conference was held at Langley Field, Virginia, on 25-26 October. Eighteen technical papers were presented to an audience of 313 individuals. Approximately ten percent of those attending the conference were representatives of various Air Force activities, and over half of these were WADC personnel. In view of the part which the Air Force had played in evaluating the original design and in the preliminary financing and procurement activities, it was surprising that there was absolutely no Air Force participation in the presentations. The majority of the twenty-seven authors who contributed papers were drawn from the NACA (16), while the remaining papers were authored by employees of the airframe (9) and engine (2) contractors.

It was evident from the papers presented at the industry conference that a considerable amount of valuable data had already been gathered but that a number of areas still awaited exploration. The airframe design differed from that originally envisaged by NACA and departed significantly from the design originally submitted by North American. The major external difference was a result of the need for additional directional stability at high angles of attack. This increased stability was provided by the addition of a ventral tail. One of the papers summarized the aerodynamic characteristics that had been obtained by tests in eight different wind tunnel facilities. c These tests had been made at Mach numbers ranging from less than 0.1 to about 6.9. The wind tunnel investigations were concerned with such problems as the effects of speed-brake deflection on drag, the lift-drag relationship of the entire aircraft, of individual components such as the wings and fairings, and of combinations of individual components. One of the interesting products was a finding that almost half of the total lift at high Mach numbers would be derived from the body-side fairing portion of the airplane. Another result was the confirmation of NACA's prediction that the original side fairings would cause longitudinal instability. (For subsequent testing the fairings had been shortened in the area ahead of the wing.) Still other wind tunnel tests had been conducted in an effort to establish the effect of the vertical and horizontal tail surfaces on longitudinal, directional, and lateral stability. Results of the wind tunnel tests were used to calculate the response characteristics of a configuration without dampers in order to determine if the aircraft would be flyable if the dampers should fail. Results indicated considerable instability and further investigations of alternate tail and rudder configurations were undertaken.

Other papers presented at the industry conference dealt with research into the effect of the aircraft's aerodynamic characteristics on the pilot's control. pilot-controlled simulation flights for the exit and entry phases had been conducted; researchers reported that the pilots had found the early configurations unflyable without damping, and that even with dampers the airplane possessed Only minimum stability for portions of the programmed flight plan. A program utilizing a free-flying model had proved low-speed stability and control to be adequate.

As some aerodynamicists had questioned North American's substitution of a differentially-operated horizontal tail for aileron control, the free-flying model had also been used to investigate that feature. The results indicated that such a tail provided the necessary lateral control.

Three of the papers presented at the conference dealt with aerodynamic heating. The first of these was a summary of the experience gained with the Bell X-1B. and X-2 aircraft. The information was incomplete and not fully applicable to the X-15, but it did provide a basis for comparison with the results of the wind tunnel and analytical studies. The second paper contained information derived from wind-tunnel tests on various bodies similar to those employed in the X-15. The third paper dealt with the results of the structural temperature estimates that had been arrived at analytically. It was apparent from the contents of the papers on aerodynamic heating that the engineers compiling them were confronted by a paradox. In order to attain an adequate and reasonably safe research vehicle, they had to foresee and compensate for the very aerodynamic heating problems that were to be explored by the completed aircraft.

In addition to the papers on the theoretical aspects of aerodynamic heating, a report was made on the structural design that had been accomplished at the time of the conference. The paper dealt with the wing, fuselage, and empennage. As critical loads would be encountered during the accelerations at launch weight and during reentry into the atmosphere, and as maximum temperatures would be encountered only during the second of these two phases, the paper was largely confined to the results of the investigations of the load-temperature relationships that were anticipated for the reentry phase. The selection of Inconel-X sheet as the covering for the multi-spar box-beam wing was justified on the basis of the strength and favorable creep characteristics of that material at 1,200 degrees Fahrenheit. A milled bar of Inconel X was to be utilized for the leading edge, as it was intended that that portion of the wing act as a heat sink. The internal structure of the wing was to be of titanium-alloy sheet and extrusions. The front and rear spars were to be flat web channel sections with the intermediate spars and ribs of corrugated webs of the same material. For purposes of the tests the maximum temperature differences between the upper and lower wing surfaces had been estimated to be 400 degrees Fahrenheit and that between the skin and the center of the spar as 960 degrees. Laboratory tests indicated that such differences could be tolerated without any adverse effects on the structure. Other tests had proved that thermal stresses for the Inconel-titanium structure were less than those encountered in similar structures constructed entirely on Inconel. Full scale tests had been made to determine the effects of temperature on the buckling and ultimate strength of a box beam, the amount of the deformations at varying loads, temperatures and temperature differences, to ascertain creep effects due to repeated loads and heating, to evaluate structural attachments and the effect of large temperature differences on the bending stresses of the spars. Simply heating the test structure produced no surface buckles. Compression buckles had appeared when ultimate loads were applied at normal temperatures but the buckles disappeared with the removal of the load. Tests at higher temperatures and involving large temperature differences had finally led to the failure of the test box, but it seemed safe to conclude that "thermal stresses had very little effect on the ultimate strength of the box."

Tests similar to those conducted on the wing structure had also been performed on the horizontal stabilizer. The planned stabilizer structure differed from the wing in that it incorporated a stainless steel spar about halfway between the leading and trailing edges, and an Inconel spar three and one-half inches from the leading edge. The remainder of the internal structure was to be similar to that of the wing in that it incorporated titanium components. The stabilizer skin was similar to that of the wing in being of Inconel-X sheet. Tests of the stabilizer had indicated that a design which would prevent all skin buckling would be inordinately heavy, so engineers decided to tolerate temporary buckles.. The proposed stabilizer had flutter characteristics that were within acceptable limits.

Brief summaries of the vertical tail and speed brake structures were also presented but as these components ultimately underwent extensive modifications, the items described had little relation to the final design.

The fuselage was to be of Inconel X. A semi-monocoque structure of titanium ribs and an inner aluminum skin were to be employed in the area ahead of the propellant tanks, and that section was to be insulated with spun glass. In the area of the propellant tanks, the circular fuselage was to be of full monocoque construction. One speaker pointed out that a full monocoque design would utilize only slightly thicker skins than a semi-monocoque design, would possess adequate heat sink properties, would reduce stresses caused by temperature differences by placing all of the material at the surface, and that the resulting structure would be ideal for use as a pressure tank. The design eliminated skin buckling and bulging, provided stiffness, had a uniformity that reduced fatigue and creep problems, and was simple to fabricate. The thickness of the monocoque walls would also make sealing easier and leaks less likely.

Fuselage problems which had not been resolved at the time of the industry conference included the reduction in buckling strength that was anticipated in the areas where the cooler internal rings of the tank bulkheads and wing support frames restrained the heated outer shell. It was known that this restraint would induce compression stresses in the shell and thereby reduce buckling strength. Another problem arose because of the side tunnels incorporated in the design. As the tunnels would protect the side portions of the circular shell from aerodynamic heating, the sides would not expand as rapidly as the areas exposed to the air and another undesirable compressive stress had to be anticipated. It was thought that beading the skin of the areas protected by the tunnels would provide a satisfactory solution but beading introduced further complications by reducing the structure's ability to carry pressure loads.

Structural design in the case of the X-15 definitely involved the propellant tanks. Each of the two main tanks was to be divided into three compartments by torus (curved) bulkheads; the two compartments furthest from the aircraft center of gravity were to be subdivided by slosh baffles. Plumbing was to be installed in a single compartment, the compartment sealed by a bulkhead, and the process repeated until all the compartments were completed. The tank ends were to be semi-torus in shape to keep them as flat as possible, to reduce weight, and to permit thermal expansion of the tank shell. This entire structure was to be of welded Inconel X. At the time of the industry conference a full-size test specimen was under construction for the purpose of testing tank pressures, external loads, temperature environments and leakage rates. A wing support frame and a section of the fuselage tunnel were to be included in the test structure in the hope that the experimental section would provide valuable static test data prior to the completion of an actual fuselage for the X-15.

Because the X-15 was expected to produce large accelerations, it seemed best to develop a side controller that would allow the pilot's arm to be restrained by an armrest without depriving him of full control over the aircraft. At the time of the industry conference in 1956, the design for the X-15 side controller had not been definitely established but a summary of the previous experience with such controllers was available. Experimental controllers had been installed on a Grumman F9F-2, a Lockheed TV-2, a Convair F-102, and on a simulator. The pilots who had tried side controllers had reported no difficulty in maneuvering, but they generally felt that greater efforts would have to be made to eliminate backlash and to control friction forces; they had also urged that efforts be made to give the side controllers a more "natural" feel.

Another problem which had not been thoroughly explored at the time of the 1956 conference concerned the proposed reaction controls that would be necessary for the X-15 as dynamic pressures decreased to the point where the aerodynamic controls would no longer be effective. Analog-computer and ground-simulator studies were then under way in an effort to determine the best relationship between the control thrust and the pilot's movement of the control stick. Attempts were also being made to determine the amount of fuel that would be required for the control rockets. No significant problems were uncovered during these early investigations, but it was clear that the pilot would have to give almost constant attention to such a control system and that pilots who were to use this form of control should be given extensive practice on simulators before being allowed to attempt actual flight.

As in the case of the other papers presented to the 1956 industry conference, the report on ground and aircraft instrumentation was very tentative in nature. Nevertheless, plans were already well along for the establishment of ground tracking stations to assist the pilot with data and advice, to record accurate measurements, and to provide navigational assistance to both the X-15 and its mother aircraft. Such a range would also prove valuable for search in case of emergency. This ground range was to be established along a line extending from Wendover AFB, Utah, to Edwards AFB, California, and was to have installations at Ely and Beatty in Nevada as well as at Edwards. The range was to be equipped so as to determine velocity, range, elevation, and azimuth with radar. Engine and aerodynamic data were to be transmitted from the X-15 by telemetering and voice radio. Each ground station was to overlap the next and all were to be interconnected so that timing signals, voice communication, and radar data would be available to all. The timing signals were to originate at Edwards. Provision was to be made for recording the acquired data on tape and film; some was to be directly displayed. Design and fabrication of this complex had been undertaken by the Electronic Engineering Company of Los Angeles. Project planners estimated that the range would be ready for operation by 1958.

In the X-15 itself, provision was being made for a pressure recorder in the nose, a main instrument compartment directly behind the pilot, and for accelerometers and other small sensing devices in a center-of-gravity compartment.

Some of the anticipated difficulties in the field of instrumentation arose because available Strain gauges were not considered satisfactory at the expected high temperatures and because of difficulties in recording the output of thermocouples. Large structural deformations of wings and empennage were to be recorded by cameras in special camera compartments.

Another instrumentation problem arose because the sensing of static pressure, ordinarily difficult at high Mach numbers, was compounded in the case of the X-15 by heating that would be too great for any conventional probe and by the low pressure at the high altitudes to be explored. Project personnel hoped that a stable-platform-integrating-accelerometer system could be developed to provide velocity, altitude, pitch, yaw, and roll angle information. Available accelerometer systems were limited to two axes and were too large and heavy for X-15 use, but it appeared that a three-axis platform within the space and weight limitations of the X-15 could be developed, and at the time of the meeting in 1956, manufacturer's proposals for such a system were being considered.

An unsolved problem was that of recording outside temperatures. The only solution appeared to be the use of radiosondes, but that was not completely satisfactory as such devices were limited to altitudes of about 100,000 feet, far less than the altitude to be attained by the X-15.

Still another instrumentation difficulty was created by the desirability of presenting the pilot with angle-of-attack and side slip information, especially for the critical exit and reentry periods. Any device to furnish this information would have to be located ahead of the aircraft's own flow disturbances, would have to be structurally sound at elevated temperatures, would have to be accurate at low pressures, and would have to cause a minimum of flow disturbance so as not to interfere with the heat transfer studies that were to be conducted in the forward area of the fuselage. These requirements had led to the development of a null balance sensing device, preliminary work had resulted in the design of a six-inch Inconel sphere capable of withstanding 1,200 degree temperatures. The sphere, to be placed in the nose of the X-15, was to be gimbaled and servo-driven in two planes. It was to have five openings: a total-head port opening directly forward and two pairs of angle-sensing ports in the pitch and yaw planes, located at an angle of 30 to 40 degrees from the central port. (Pitch and yaw of an aircraft could be sensed as pressure differences and these differences converted into signals that would cause the servos to realign the sphere in the relative wind.) As a null-balance device had no source of static pressure, it was not suitable for furnishing indicated airspeed, so some alternate pitot-static system would be necessary to provide the airspeed information required for landing the X-15 safely.

Two main criteria had governed the selection of an escape system for the X-15, and these two criteria were not necessarily complementary. The first requirement had been that the system be the most suitable that could be designed while remaining compatible with the airplane. The second requirement had been that no system would be selected that would delay the development of the X-15 or leave the pilot without any method of escape when the time arrived for flight testing the completed vehicle. The four possible escape systems that were considered included cockpit capsules, nose capsules, a canopy shielded seat, and a stable-seat, pressure-suit combination. An analysis of the expected flight hazards had indicated that because of the fuel exhaustion and low aerodynamic loads, the accident potential at peak speeds and altitudes was only about two percent of the total accident potential.

The final decision for a stable-seat, pressure-suit combination was made because most of the potential accidents could be expected to occur at speeds of Mach 4 or less, because system reliability always decreased with system complexity, and finally, because it was the system that imposed the smallest weight and size penalties upon the aircraft. The selected system would not function successfully at altitudes above 120,000 feet and speeds in excess of Mach 4, but designers held that the aircraft itself would be its own best escape capsule in the areas where the seat-suit combination was inadequate.

The preliminary ejection seat design utilized a rocket-type ejection gun. One proposed version was fin-stabilized and another incorporated a skip-flow generator. d A preliminary decision had been made to use the skip-flow type. The seat also incorporated restraining devices for the pilot's extremities.

An emergency oxygen system was to be capable of providing suit pressurization and a breathing supply for a period of twenty minutes. The pressure suit was to be similar to those already in development for high performance military aircraft. Such a suit was considered adequate for protection against the ozone hazard and it had been decided that there was no necessity for concern with exposure to cosmic rays. Concern was expressed, however, for the problem of rapid pressure changes during the various stages of the ejection sequence. Researchers concluded that careful consideration would have to be given to the possible pressure surges within the helmet and their potential for damaging the pilot's ears and lungs. It had already been determined that the proposed suit materials could withstand the maximum pressure and temperatures to which they would be subjected within the operational limits of the escape system as a whole.

The plans for the cockpit environment of the X-15 were based on the use of nitrogen. Cockpit and instrument cooling, pressurization, suit ventilation, windshield defogging, and fire protection were all to be provided from a liquid nitrogen supply. Vaporization of the liquid nitrogen would keep the pilot's environment within comfortable limits at all times. An interesting aspect of the cooling problem was an estimate that only 1.5 percent of the system's capacity would be applied to the pilot; the remaining 98.5 percent was required for the equipment. Cockpit temperatures were to be limited to no more than 150 degrees, the maximum limit for some of the equipment. The pilot would not be subjected to that temperature, however, as the pressure suit ventilation would enable him to select a comfortable temperature level for himself. Cockpit pressure was to be maintained at the 35,000 foot level and as the pressure suit was also designed to operate at the same level if cabin pressure should fail, there would be no pressure variations during the exploratory phases of the X-15's flights and the pilot would have adequate protection against explosive decompression. Provision was to be made for the pilot to clear the cockpit area of its nitrogen atmosphere by the use of ram air pressure.

The various switches and controls were to be selected and placed to minimize pilot movements. Instrument, warning light, and control location had been determined by analysis of the pilot's duties and instruments were to be arranged in a manner that would permit a maximum of attention to be directed toward one area at a time. Visibility from the cockpit would be excellent, but some questions remained unanswered as to the vision-degrading effects of heat distortion from the hot windshield. Key to detailed layout of the cockpit was the planned use of side controllers and the possible elimination of the center stick. Scott Crossfield, a former NACA test pilot who was to make the initial X-15 flights as a North American pilot, commented that the decision to abandon the center stick would rest on the results of further tests and the necessity to "break with tradition." (He may have reflected that the world's first military airplane was guided with side controllers and that it had no center stick.)

The effects of flight accelerations upon the pilot's physiological condition and upon his ability to avoid inadvertent control movements had not been completely explored, but it was recognized that high accelerations could pose medical and restraint difficulties. In addition to the accelerations that would be encountered during the exit and reentry phases of the X-15's flights, a very high acceleration of short duration would be produced during the landings. This latter acceleration was a result of the location of the main skids at the rear of the aircraft. Once the skids touched down, the entire aircraft would act as if it were hinged at the skid attachment points and the nose section would slam downward. Reproduction of this landing acceleration on simulators showed that because of the short duration, no real problem existed. There were however, numerous complaints about the severity of the jolts.

The 1956 industry conference heard two papers on the proposed engine and propulsion system for the X-15. The first of these dealt only with the engine, the second with the installation of the engine and its associated systems in the aircraft. At the time of the conference the proposed XLR99-RM-1 engine was scheduled to have a variable thrust of from 19,200 to 57,200 pounds at 40,000 feet. It was to employ anhydrous ammonia, liquid oxygen and a 90-percent hydrogen peroxide solution as propellants, was to have a dry weight of 618 pounds, and a wet weight of 748 pounds. Specific impulse was to vary from a minimum of 256 seconds to a maximum of 276 seconds. The proposed engine was to fit into a space with a length of 71.7 inches and a diameter of 43.2 inches. A single thrust chamber was to be supplied by a turbopump with the turbopump's exhaust being recovered in the thrust chamber. A two-stage impulse turbine was to drive a dual inlet fuel pump and a single inlet oxidizer pump. Thrust control was by regulation of the turbopump speed, the regulation to be accomplished by a pilot-controlled governor. e

In the design stages of the XLR99's development, Reaction Motors was concerned with the engine's safety and reliability in terms of the requirement to produce an engine that could be throttled and that would meet the established specifications. (Another factor of some importance was the requirement that the engine should be capable of being restarted.)

The decision to control the engine's thrust by regulation of the turbopump's speed was made because the other possibilities (regulation by measurement of the pressure in the thrust chamber or of the pressure of the discharge) would cause the turbopump to speed up as pressure dropped. As the most likely cause of pressure drop would be cavitation in the propellant system, an increase in turbopump speed would aggravate rather than correct the situation. Reaction Motors had also decided that varying the injection area was too complicated a method for attaining a variable thrust engine and had chosen to vary the injection pressure instead.

The regenerative cooling of the thrust chamber created another problem for the designers as the varying fuel flow of a throttleable engine meant that the system's cooling capacity would also vary and that adequate cooling throughout the engine's operating range would produce excess cooling under some conditions. Engine compartment temperatures also had to be given more consideration than in previous rocket engine designs because of the higher radiant heat transfer from the structure of the X-15.

The restart requirements for the XLR99 introduced some additional complications, particularly in regard to safety provisions. At the time of the conference, a two-stage ignition system was planned; the effort to produce a fail-safe design for the ignition system and the engine itself necessitated a purge system, inert gas bleed for both stages of the ignition and thrust chamber, and the duplication of numerous system components. On the other hand, the fact that both fuel and oxidizer were volatile reduced the hazard of an unsafe accumulation of propellants in the system.

Reaction Motor's spokesman at the conference of 1956 concluded that the development of the XLR99 was going to be a difficult task. Subsequent events were certainly to prove the validity of that assumption.

A second paper dealt with engine and accessory installation, the location of the propellant system components, and the engine controls and instruments. The main propellant tanks were to contain the liquid oxygen (LOX), ammonia, and the hydrogen peroxide. The oxygen tank, with a capacity of approximately 1,000 gallons, was to be located just ahead of the aircraft's center of gravity; the ammonia tank, with a capacity of approximately 1,400 gallons, just aft of the same point. A center core tube within the oxygen tank would provide a location for a supply of helium under a pressure of 3,600 pounds per square inch. Helium was to be utilized for the pressurization of both the oxygen and ammonia tanks. A 75-gallon hydrogen peroxide tank behind the ammonia tank was to provide the monopropellant for the engine's turbopump. An additional supply of helium was to be utilized for pressurizing the monopropellant tank. The LOX and ammonia tanks were designed with triple compartments arranged to permit both propellants to be forced toward the center of gravity as they were expelled, either during normal operations or jettisoning. The transfer tubes between compartments demanded considerable study because the high accelerations of the X-15 would tend to force the contents of the tanks toward one end or the other. The compartmental divisions were further complicated by the necessity for efficient fueling and the need to keep the quantity of propellants remaining after burnout or jettisoning at the lowest possible figure. As the acceleration, efficient fueling, and maximum evacuation called for features not entirely compatible, compromises were necessary. Fortunately, no insoluble problems arose during early tests.

Provision was also made for top-off of the LOX tank from a supply carried aloft by the mother aircraft. Top-off from the mother airplane was considered to be beneficial in two ways. The LOX supply in the mother ship could be kept cooler than the oxygen already aboard the X-15, and the added LOX would permit cooling of the X-15's own supply by boil-off, without reduction of the quantity available for flight. The ammonia tank was not to be provided with a top-off arrangement, as the slight increase in fuel temperature during carried flight was not considered significant enough to justify the complications such a system would have entailed.

A suitable material had not yet been selected for the tank that was to contain the high-pressure, low-temperature helium supply for propellant tank pressurization. The entire propellant system presented problems difficult to foresee, primarily because of the large variations of temperature and pressure that would occur during a single flight of the X-15.

Because engine vibration characteristics were unknown, the engine mount was designed to be rigid without any special effort at vibration shielding. The engine-mount truss was to join the thrust chamber at several points and was to be attached to the fuselage by three fittings designed so that the top attachment provided the main pivot point. The two lower fittings were to be adjustable to allow accurate alignment of the engine's thrust vector.

Three large removable doors were to provide access to the engine area and to permit observation of the engine by closed circuit television cameras during ground testing of the engine. The entire engine compartment was designed to explode open at a pressure lower than that which the forward structure was capable of withstanding, thus providing relief in case of an engine explosion. As engine compartment temperatures were not expected to be a problem, no insulation was being planned.

In 1956, the cockpit engine instruments had not been finally selected and the preliminary choices were to be altered as the engine and the propellant system were developed. A throttle with an engine prime switch was to be located on the left console, with the tank pressurization switches and jettisoning controls in the immediate vicinity of the throttle. Electric switches were to be provided for engine arming, for fire extinguisher control, and for master control. Space was allotted for several indicators to furnish the pilot with pressure information on the propellant and engine systems. A place had also been reserved for six lights to indicate various engine malfunctions. It had also been decided that the pilot would need an instrument (a totalizing impulse indicator) capable of showing the total thrust remaining at any given instant during powered flight.

The final paper presented to the 1956 industry conference was a summary of the preceding papers and of the major problems that existed at that time. The author considered flutter to be an unsolved problem, primarily because of a lack of basic data on aero-thermal-elastic relationships and because little experimental data was available on flutter at the hypersonic Mach numbers that would be reached by the X-15. He pointed out that available data on high speed flutter had been derived from experiments conducted at Mach 3 or less, and that not all of the data obtained at those speeds were applicable to the problems faced by the designers of the X-15. He felt that the solution of the problem was full-scale robot testing of X-15 components. Another difficulty was the newness of Inconel-X as a structural material and the necessity of experimenting with fabrication techniques that would permit its use as the primary structural material for the X-15. Problems were also expected to arise in connection with sealing materials, most of which were known to react unfavorably when subjected to high temperature conditions. Preliminary wind-tunnel tests had also indicated that the original configuration of the X-15 did not have adequate stability and that modification and further testing would be essential.

The closing portion of the final paper dealt briefly with North American's schedule for drawings, jig construction, and fabrication of the aircraft. 8

That the design features of the X-15 presented to the industry conference in October 1956 were only tentative was made apparent by the results of a development engineering inspection held at North American's Inglewood plant on 12-13 December 1956. This inspection of a full-scale mockup was intended to reveal unsatisfactory design features before fabrication of the aircraft got under way. Thirty-four of the forty-nine individuals who participated in the inspection were representatives of the Air Force, and twenty-two of them were from Wright Air Development Center. The important role of the Air Force in the determination of the X-15's design was evident from the composition of the committee chosen to review the alteration requests. Major E. C. Freeman, of the Air Research and Development Command, served as committee chairman, Mr. F. Orazio of Wright Air Development Center and Lieutenant Colonel K. C. Lindell of Air Force headquarters were committee members, and Captain C. E. McCollough Jr. of the Air Research and Development Command and Captain I. C. Kincheloe of the Air Force Flight Test Center served as advisors. The Navy and NACA each provided a single committee member; three additional advisors were drawn from the staff of the NACA.

The inspection committee considered 84 requests for alterations, decided to reject 12, and placed 22 in a category requiring further study. The change requests covered a variety of features, including the controls, electrical and hydraulic systems, the escape system, and the power plant. Some of the accepted changes were the addition of longitudinal trim indications from the stick position and trim switches, relocation of the battery switch, removal of landing gear warning lights, rearrangement and redesign of warning lights, and improved marking for several instruments and controls. Other accepted recommendations concerned improved wiring for the fire detection system, improved insulation of sensitive electrical equipment, inclusion of an overheat warning system for hydrogen peroxide compartments, and the relocation of some of the electrical wiring in order to protect it from hydraulic fluids and to reduce the possibility of damage during the installation and removal of equipment. Inspection personnel also requested that the escape system be provided with better markings, that safety pins be identified by streamers, and that a dependable linkage be installed between the canopy and seat catapult initiator. Still other approved changes concerned such items as a lock for the LOX filler cover, and improvement of the hydraulic system by the substitution of some components and by better installation and marking. The landing gear was the subject of a number of suggestions, including the elimination of cadmium plating on certain heat treated steels employed in the gear, provision for inspection panels, the use of new tires on each of the early flights, and for additional design and testing of all components of the skids and nose gear.

The requested changes in the propulsion system were concerned with the improvement of the hydrogen peroxide system by the inclusion of better leak protection methods, by better support for the tanks, and relocation of shut-off valves. Inspection personnel also recommended attention be given to keeping engine components in locations where they could be easily inspected and maintained, that adequate drainage and ventilation be provided for the engine compartment, that North American provide engine mounts to Reaction Motors in order to simplify engine handling and installation, and that engine mount bolts be safetied. Improvements were also asked if the design of the jettisoning system and in the identification and marking of the propellant system's components.

Some of the most interesting of the proposed changes were rejected by the committee. For instance, the suggestions that the aerodynamic and reaction controller motions be made similar, that the reaction controls be made operable by the same controller utilized for the aerodynamic controls, or that a third controller combining the functions of the aerodynamic and reaction controllers be added to the right console, were all rejected on the grounds that actual flight experience was needed with the controllers already selected before a decision could be made on worthwhile improvements or combinations. As two of the three suggestions on the controllers came from potential pilots of the X-15 (J. A. Walker of the NACA and Captain Kincheloe), it would appear that the planned controllers were not all that might have been desired. A warning light for the canopy lock was also rejected, as was the suggestion that the pilot be provided with easier entrance and exit by extension of the canopy's travel - both on the grounds that the existing provisions were adequate. Simplification of the hydraulic system on the first airplane was ruled out on the basis that there was nothing that could be spared. A request that the pilot be provided with continuous information on the nose-wheel door position (because loss of the door could produce severe structural damage) was rejected because the committee felt that the previously approved suggestion for gear-up inspection panels would make such information unnecessary. A suggested study of the ignition and fire hazard potential of the various mixtures that might accumulate in the engine compartment was held to be unnecessary in light of the ventilation provisions for that compartment. Joining the auxiliary power plant exhausts in a single manifold to avoid out-of-trim moments if one auxiliary power plant should fail was not felt to be necessary. A suggested addition of check valves in the hydrogen peroxide system was considered to have been adequately taken care of by previously accepted suggestions. A request for entirely separate systems for each auxiliary power unit and for the ballistic controls was supported by the argument that separate systems had been requested earlier. (Such separate systems had been accepted during the meetings held at the North American plant in the fall of 1955.) In spite of the earlier plans for such separate systems, the committee held that with the addition of shut off valves, the system would be adequate as installed.

An even more surprising rejection of a requested change occurred in regard to changeable leading edges. An NACA representative (Harry J. Goett of Ames Aeronautical Laboratory) asked that the lower flange of the front spar be widened and that the ballistic roll controls be moved to the rear of the same spar. He justified these requests on the grounds that the research goals for the X-15 included investigations to determine the best materials, profiles, and cooling methods for various leading edges; that interchangeable leading edges had been a part of the original proposals; and that North American had originally agreed to make the leading edge detachable. In spite of Mr. Goett's apparently logical arguments, the committee decided his request could not be honored. The reasons for their rejection of the request were that North American had already determined to use a solid plate for the lower wing surface and that the required changes would impose a three-pound weight penalty. It seemed to at least one participant that the negative decision on interchangeable leading edges marked the abandonment of a feature that would have considerably enhanced the research value of the X-15.

That a number of design features still were unsettled as late as the mockup inspection of 13 December 1956, was indicated by the 22 change requests placed in a category requiring further study. Some of these deferred requests were concerned with the B-36 carrier aircraft, which was eventually eliminated; other change requests required feasibility studies, however, and some needed further study as to desirability. The deferred requests included such suggestions as the complete elimination of the center control stick, a study of antenna locations to insure there would be no adverse effects on directional stability, the installation of an engine ignition gauge, a LOX top-off indicator, and improvements in the rigidity and alignment of the accelerometer mounts. Doubts were expressed about the adequacy of a single antenna for transmitting and receiving radar signals and further studies were promised. A request for hydraulic pressure indication when both generators were out was also deferred until it could be determined if such indication was feasible. Three deferred requests on the escape system involved the continuing development of seat and pressure suit by further sled and tunnel tests, a study to determine the desirability of a spoiler plate to be located ahead of the cockpit and operated in the canopy ejection sequence, and the selection of an improved location for the canopy's emergency release handle.

Other requests which the committee decided to be worth further study included the replacement of machine screws by quick fasteners for some of the fuselage access panels, vibration testing of propellant lines, relocation of components of the helium system to minimize the possibility of leaks, and the use of expulsion bags in the hydrogen peroxide tanks. Further study was to be conducted to determine the best type of bag or diaphragm for hydrogen peroxide expulsion, the adequacy of the helium tank mounts, the ability of the propellant lines to withstand stresses imposed by engine misalignment, and the feasibility of starting the engine ignition system prior to launch. A request to approve the shift of all controls and switches to locations where they could be easily reached from the pilot's normal seated position, (even when the pilot was of small stature), received the "further study" classification, but in this case the group also authorized such changes as appeared necessary. 9

After the completion of the development engineering inspection, the X-15 airframe design changed only in relatively minor details. North American essentially built the X-15 described at the industry conference in October and inspected in mockup in December. (Continued wind tunnel testing resulted in some external modifications, particularly of the vertical tail, and some weight changes occurred as plans became more definite.) But while work on the airframe progressed smoothly, with few unexpected problems, the project as a whole did encounter difficulties, some of them serious enough to threaten long delays. In fact, North American's rapid preparation of drawings and production planning served to highlight the lack of progress on some of the components and sub-systems that were essential to the success of the program.